1. Field of Disclosure
The disclosure relates generally to an onboard navigation system and more particularly, to a system and method for estimating the bias errors associated with inertial acceleration measurements for precision guidance of vehicles.
2. Related Art
The aircraft industry has developed automatic landing capability using a differential Global Positioning System (GPS). This capability is known as the Global Navigation Satellite System Landing System (GLS). The GLS has been certified for CAT operations and the aircraft industry is now developing standards and performance requirements for GLS to support CAT II/III operations.
The GLS includes an onboard inertial sensor suite having accelerometers and angular rate sensors (gyroscopes), augmented by position and/or velocity updates derived from external measuring systems, such as GPS, used to track a vehicle's movements and position. The onboard sensors typically have sufficient accuracy to maintain a useable navigation solution during the small periods of time in between the external updates.
However, a key issue associated with GLS CAT II/III operations are the expected GLS failure modes and their affect on the aircraft guidance system. It is anticipated that the most common failure mode for GLS will be a total loss of external updates, such as the GLS positions and velocities, for an extended period of time, such as for hundreds of seconds. As a result, the loss of external updates causes the location of the vehicle to be inaccurate. If this failure were to occur below a specified alert height, the system must tolerate the failure and continue to perform an acceptable autoland.
To ensure an acceptable autoland, a GPS/INS filtering scheme was developed and is the subject of U.S. Pat. No. 6,178,363 (U.S. Pat. No. '363) entitled “Inertially Augmented GPS Landing System”, issued on Jan. 23, 2001, which is assigned to the assignee of the present application, and the disclosure of which is incorporated herein by reference in its entirety. The filtering scheme disclosed in U.S. Pat. No. '363 smoothes the GLS deviations from runway centerline and glidepath with inertial velocities and/or accelerations.
The resulting filtered deviation outputs can sustain an interruption of the GLS position and velocity updates through the double integration of the inertial acceleration measurements summed with an inertial bias error correction term. However, this initial filtering technique relies on instantaneous detection of a GLS system failure. If there is any delay in the detection of the GLS failure, the inertial bias correction term is subject to corruption by erroneous GLS position information during the period of time that the GLS failure is present but undetected. To overcome the potential corruption due to a failing GLS signal, the structure of the complementary filter may be modified and combined with an integrator reset scheme which is the subject of U.S. Pat. No. 6,549,829 (U.S. Pat. No. '829) entitled “Skipping Filter For Inertially Augmented Landing System”, issued on Apr. 15, 2003, which is assigned to the assignee of the present application, and the disclosure of which is incorporated herein by reference in its entirety. By resetting the filter using buffered values which are guaranteed not to have been corrupted, and removing the GLS input from the filter to eliminate future corruption, the filter can continue to provide guidance based strictly on the corrected inertial information.
In an attempt to decrease the cost of aircraft, it is becoming more common to use less expensive (and less accurate) inertial sensors. Overall navigation system accuracy may still be satisfactory because GPS position information is now readily available to update the navigation solution. Unfortunately this poses a problem for GLS approaches because the same position information being sent to the inertial units for synthesis of the GPS blended inertial navigation solution is also being used to synthesize the GLS deviations. If no alternative inertial outputs (independent of GPS/GLS failures) are available, fail-operative autoland in the presence of GLS signal loss is compromised by potential corruption from the GPS supplement to the inertial rates. This corruption, entering the filter via the inertial data, can not be removed as the inertial data must be independent of the GLS data to remove corruption entering the filter via the failing GLS input.
To eliminate the potential for corruption of the filter via the inertial data input, the uncorrected acceleration measurements may be transformed into runway coordinates and blended with the GLS deviations. Both the complementary filtering schemes in U.S. Pat. No. '363 and U.S. Pat. No. '829 depend on a common reference frame between the inertial accelerations and/or inertial velocities and the independent position and/or velocity source update. However, a pure translation of accelerations is not enough to guarantee accurate bias estimation, depending on the behavior of the inertial error. For example, if the inertial accelerations and/or velocities being used in the filter are output as part of the inertial reference unites (IRU's) navigation solution, the filtering within the IRU fixes the dominant inertial errors (Schuler errors) to the navigation-frame (coincident with the GLS reference frame). Thus the same cross-runway bias is detected by the inertial/GLS comparison within the filter regardless of the aircraft's attitudes with respect to the runway.
Alternatively if the inertial accelerations and/or inertial velocities available for use in such a filter are not pre-processed, as in this example where the only inertial data which cannot be corrupted by GPS failures are the uncorrected accelerometer outputs, the dominant inertial biases are fixed to the accelerometer orientation, regardless of the transformation of the accelerations themselves. This results in a mismatch in the comparison of the inertial error (fixed to the local accelerometer frame) and the fixed-earth deviations referenced to the runway which causes the cross-runway bias detectable by the previously patented filter schemes to be subject to any differences between the aircraft attitude and the runway.
With a bias estimate which is subject to aircraft attitude changes the bias perceived on approach (including a crab angle for example) may be significantly different than that perceived during rollout (where the crab angle has been removed) due to the difference in along and across body acceleration components synthesizing the cross-runway acceleration. Thus, the bias estimated through the approach is potentially non applicable when the GLS signal becomes invalid. In this case, without an alternative means of estimating the inertial errors in their local frame, a potentially costly upgrade in the IRU is required to outfit it with either high accuracy gyros or additional computational power to handle the added complexity of a separate navigation solution output synthesized without GPS aiding to enable fail-operational GLS autoland.
In view of the above, what is needed is an alternative system and method of estimating inertial errors in their local frames.